A cast heat exchanger, having an inlet and an outlet and defining an internal passage therebetween, includes a wall structure disposed within the internal passage that defines a series of interconnected chambers through which a cooling medium flows in a serpentine path. The outer surface of the heat exchanger is covered with pin fins that transfer heat away from the hot working medium flowing over the pin fins and into the cooling medium within the chambers of the heat exchanger.
|
1. An apparatus comprising:
a gas turbine engine having a compressor section and a combustor section; a cast housing located within said combustor section and having an outer surface within and in a heat transfer relationship with a flow of a compressed air from said compressor section, said cast housing having an inlet and an outlet, said cast housing defining an internal passage between said inlet and outlet; and said housing further including a wall structure disposed within said internal passage to define a serpentine path from said inlet to said outlet.
11. An apparatus comprising:
a gas turbine engine having a compressor portion, a combustor portion and a turbine portion; a cast heat exchanger within said gas turbine engine and having an outer surface disposed within and in a heat transfer relationship with a flow of air from said compressor portion, said heat exchanger having a cooling media inlet and a cooling media outlet, and defining therebetween an internal passage adapted to contain a cooling media; and said heat exchanger further including means for flowing said cooling media in a serpentine path through said internal passage.
19. An apparatus, comprising:
a gas turbine engine having a compressor and combustor; a plurality of cast heat exchangers placed in series axially within said gas turbine engine and disposed within a flow of air from said compressor, each heat exchanger including an inlet and an outlet and defining an internal passage therebetween, each heat exchanger further including means for flowing a cooling medium in a serpentine path through said internal passage, and wherein said internal passages are coupled in fluid communication with one another to define a pathway within said plurality of heat exchangers.
2. The apparatus of
3. The apparatus of
5. The apparatus of
6. The apparatus of
7. The apparatus of
9. The apparatus of
10. The apparatus of
wherein said housing has a partial annular shape and said outer surface has a plurality of heat transfer members integrally formed therewith; wherein said wall structure comprises a circumferential wall and a plurality of transversely extending walls to define a plurality of chambers disposed in flow communication along said serpentine path; and wherein said cast housing and said wall structure are integrally cast of a high heat transfer coefficient metallic material.
12. The apparatus of
13. The apparatus of
15. The apparatus of
16. The apparatus of
17. The apparatus of
18. The apparatus of
20. The apparatus of
21. The apparatus of
|
This invention was made with U.S. Government support under contract number F33615-94-C-2482 and the U.S. Government may have certain rights in the invention.
This invention relates generally to heat exchangers placed within gas turbine engines. More particularly, the present invention relates to a heat exchanger having an internal passageway and being disposed within the fluid flow path prior to the combustor. Although the present invention was developed for use in a gas turbine engine, certain applications may be outside of this field.
In gas turbine engines, cooling air is generally bled off at various stages within the compressor and used for cooling elsewhere in the engine. As pressures and temperatures increase within gas turbine compressors, the temperature of the cooling air increases to a point where its usefulness as a cooling agent becomes minimal. Heat exchangers located outside the flow path of the gas turbine engine require complex piping and, therefore, introduce additional weight as well as pressure loses inefficiencies. Some studies placing heat exchangers within gas turbine engines have been undertaken and show promising results.
In
The invention described herein provides cooling means for reducing the cooling air bled off from the compressor of a gas turbine engine.
One form of the present invention contemplates a heat exchanger having an internal passageway and being disposed within the fluid flow path prior to the combustor so that the cooling air from the compressor is cooled as it flows over the outer surface of the heat exchanger.
In another embodiment of the invention, the heat exchanger is of an integral cast configuration and has a serpentine internal passageway. Eliminating the need for multiple-braze joint reduces the cost of such a cast heat exchanger.
The cooling medium flowing within the internal passageway of the heat exchanger can be fuel which cools the compressed air that is generally bled off from the compressor. The vaporized fuel is then supplied to the combustor of the gas turbine engine.
Furthermore, a series of heat exchanger segments can be arranged axially, thereby eliminating the need for separate individual heat exchangers of varying lengths.
One object of the present invention is to provide a unique heat exchanger.
Related objects and advantages of the present invention will be apparent from the following description.
For the purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiments illustrated in the drawings and specific language will be used to describe the same. It will nevertheless be understood that no limitation of the scope of the invention is thereby intended, such alterations and further modifications in the illustrated device, and such further applications of the principles of the invention as illustrated therein being contemplated as would normally occur to one skilled in the art to which the invention relates.
Referring to
A gas turbine engine is equally suited to be used for an industrial application. Historically, there has been widespread application of industrial gas turbine engines, such as pumping sets for gas and oil transmission lines, electricity generation, and naval propulsion.
The compressor section 22 includes a rotor 19 having a plurality of compressor blades 28 coupled thereto. The rotor 19 is affixed to a shaft 25 that is rotatable within the gas turbine engine 20. A plurality of compressor vanes 29 are positioned within the compressor section 22 to direct the fluid flow relative to blades 28. Turbine section 24 includes a plurality of turbine blade 30 that are coupled to rotor disk 31. The rotor disk 31 is affixed to shaft 25, which is rotatable within the gas turbine engine 20. Energy extracted in the turbine section 24 form the hot gas exiting the combustor section 23 is transmitted through shaft 25 to drive the compressor section 22. Further, a plurality of turbine vanes 32 are positioned within the turbine section 24 to direct the hot gaseous flow stream exiting the combustor section 23.
The turbine section 24 provides power to a fan shaft 26, which drives the fan section 21. The fan section 21 includes a fan 18 having a plurality of fan blades 33. Air enters the gas turbine engine 20 in the direction of arrows A and passes through the fan section 21 into the compressor section 22 and a bypass duct 27. Further details related to the principles and components of a conventional gas turbine engine will not be described herein as they are believed known to one of ordinary skill in the art.
In
Heat exchanger 40 is preferably a single cast structure having inlet port 42, outlet port 44, and internal passage 43 located therebetween. Cooling medium 41 enters heat exchanger 40 through inlet port 42, flows through internal passage 43, and exits heat exchanger 40 through outlet port 44. The cooling medium may be any suitable fluid, but in the preferred embodiment, the cooling medium is fuel. The cold liquefied fuel is heated in the process of cooling the hot compressed air, and the resulting high temperature fuel is then supplied to the combustor of the gas turbine engine. The ports 42 and 44 are designed to have a fluid flow passageway coupled thereto, and in one embodiment are tapped for a threaded fitting and in another embodiment are prepared to have fittings brazed thereto. Also, ports 42 and 44 may be laid over at different angles as required for appropriate fuel line connections to the gas turbine engine.
Heat exchanger segment 40 includes front wall 47, back wall 48, and sides 49 connecting front and back walls 47 and 48. The inner surfaces of walls 47 and 48 and of sides 49 define internal passage 43. In one embodiment, the outer surface of sides 49 is covered with fins 45. Although fins 45 may be any type of fin configuration that transfers heat away from hot compressed air flow 46 and into cool fuel 41, pin fins, as illustrated in
Disposed within internal passage 43 of heat exchanger 40 is internal wall structure 50 which comprises circumferential wall segment 51 and a plurality of radial wall segments 52. Circumferential wall segment 51 connects to front wall 47, but not back wall 48, of heat exchanger segment 40. Radial wall segments 52 extend radially from circumferential wall segment 51 and connect to a side 49 as well as back wall 48. The plurality of radial wall segments 52 further comprise a plurality of high walls 53 that connect to front wall 47 and a plurality of low walls 54 that terminate short of connecting to front wall 47.
Internal wall structure 50 thus creates a series of interconnected chambers 60, 61, 62, 63, 64, and 65 within internal passage 43 that route the fuel from inlet port 42 along serpentine path 55 to outlet port 44. In the preferred embodiment, serpentine path 55 makes multiple traverses of internal passage 43 of heat exchanger 40 in the axial direction (i.e., between front and back walls 47 and 48) and in the radial direction (i.e., between sides 49). Serpentine path 55 also circumferentially traverses internal passage 43 of heat exchanger 40. However, heat exchangers having other internal flow path configurations are contemplated herein.
Heat exchanger 40 can, but need not necessarily, be a complete 360 degree annular ring; instead, it can be a portion of a full ring unit. Multiple heat exchanger segments 40 can be located about the centerline of a gas turbine engine to yield a full ring unit if desired. Since heat exchanger 40 is cooled by fuel flow within internal passage 43, it need not necessarily be made from a high temperature alloy. It is preferably made from a material that has a high heat transfer coefficient such as beryllium copper or aluminum.
A series of exchanger segments 40 can also be placed in series axially as illustrated in FIG. 4. Air flows along path 70 over the series of heat exchangers 40. The fluid inlets and outlets of heat exchangers 40 are rolled over for a more compact design and connected to inlet manifold 72 and outlet manifold 74, respectively. Although
While the invention has been illustrated and described in detail in the drawings and foregoing description, the same is to be considered as illustrative and not restrictive in character, it being understood that only the preferred embodiment has been shown and described and that all changes and modifications that come within the spirit of the invention are desired to be protected.
Patent | Priority | Assignee | Title |
10107200, | Apr 30 2015 | General Electric Company | Turbine engine thermal management |
10626798, | Dec 09 2015 | RTX CORPORATION | Diffuser mounted fuel-air heat exchanger |
10934939, | Apr 30 2015 | General Electric Company | Turbine engine thermal management |
11209224, | Apr 19 2018 | RTX CORPORATION | Mixing between flow channels of cast plate heat exchanger |
11454169, | Dec 28 2015 | General Electric Company | Method and system for a combined air-oil cooler and fuel-oil cooler heat exchanger |
11591964, | Jun 17 2021 | Pratt & Whitney Canada Corp | Oil cooling system for aircraft engine |
7784528, | Dec 27 2006 | General Electric Company | Heat exchanger system having manifolds structurally integrated with a duct |
8601791, | Dec 23 2009 | Techspace Aero S.A. | Integration of a surface heat exchanger to the wall of an aerodynamic flowpath by a structure of reinforcement rods |
9512780, | Jul 31 2013 | GE INFRASTRUCTURE TECHNOLOGY LLC | Heat transfer assembly and methods of assembling the same |
9771867, | Dec 30 2011 | ROLLS-ROYCE NORTH AMERICAN TECHNOLOGIES INC | Gas turbine engine with air/fuel heat exchanger |
Patent | Priority | Assignee | Title |
2717581, | |||
2907170, | |||
3134536, | |||
3733816, | |||
3734639, | |||
3800868, | |||
3874168, | |||
4012912, | Apr 09 1975 | Turbine | |
5123242, | Jul 30 1990 | General Electric Company | Precooling heat exchange arrangement integral with mounting structure fairing of gas turbine engine |
5203163, | Aug 01 1990 | General Electric Company | Heat exchange arrangement in a gas turbine engine fan duct for cooling hot bleed air |
5697208, | Jun 02 1995 | Solar Turbines Incorporated | Turbine cooling cycle |
5722241, | Feb 26 1996 | SIEMENS ENERGY, INC | Integrally intercooled axial compressor and its application to power plants |
EP547641, |
Executed on | Assignor | Assignee | Conveyance | Frame | Reel | Doc |
Mar 13 2000 | Allison Advanced Development Company | (assignment on the face of the patent) | / | |||
Jul 05 2000 | RICE, EDWARD C | Allison Advanced Development Company | ASSIGNMENT OF ASSIGNORS INTEREST SEE DOCUMENT FOR DETAILS | 011029 | /0039 |
Date | Maintenance Fee Events |
Dec 14 2005 | M1551: Payment of Maintenance Fee, 4th Year, Large Entity. |
Dec 21 2005 | ASPN: Payor Number Assigned. |
Oct 20 2009 | ASPN: Payor Number Assigned. |
Oct 20 2009 | RMPN: Payer Number De-assigned. |
Jan 15 2010 | M1552: Payment of Maintenance Fee, 8th Year, Large Entity. |
Jan 17 2014 | M1553: Payment of Maintenance Fee, 12th Year, Large Entity. |
Date | Maintenance Schedule |
Jul 23 2005 | 4 years fee payment window open |
Jan 23 2006 | 6 months grace period start (w surcharge) |
Jul 23 2006 | patent expiry (for year 4) |
Jul 23 2008 | 2 years to revive unintentionally abandoned end. (for year 4) |
Jul 23 2009 | 8 years fee payment window open |
Jan 23 2010 | 6 months grace period start (w surcharge) |
Jul 23 2010 | patent expiry (for year 8) |
Jul 23 2012 | 2 years to revive unintentionally abandoned end. (for year 8) |
Jul 23 2013 | 12 years fee payment window open |
Jan 23 2014 | 6 months grace period start (w surcharge) |
Jul 23 2014 | patent expiry (for year 12) |
Jul 23 2016 | 2 years to revive unintentionally abandoned end. (for year 12) |